This application claims priority to German Patent Application DE10162238.4, filed Dec. 18, 2001, the entirety of which is incorporated by reference herein.
This invention relates to an air intake system of a propeller-turbine engine with an in-line arrangement of the propeller gearbox.
Various designs of propeller-turbine engines are known from the state of the art. These engines include a gas turbine and an upstream gearbox through which a propeller is driven.
The known designs are disadvantageous in that the size of the propeller gearbox and the bulkiness of the engine mounts typical of propeller-turbine engines necessitate a very large engine nacelle. The size of the engine nacelle is also influenced by the space required by the heat exchangers. In summary, the aerodynamic fairing of the propeller-turbine engine leads to a comparatively oversized nacelle.
A nacelle of such size has a correspondingly large frontal area, which, in turn, entails high aerodynamic nacelle drag. It should be noted in this context that propeller-turbine engines have a smaller airflow requirement, this resulting in low air throughput of the nacelle.
A further disadvantage of the known designs lies in the fact that the air intake ducts required have a relatively complicated shape. This results in a non-uniform field of flow in the air intake of the compressor, which, in turn, may become the main cause of compressor instability.
It should further be noted that the aerodynamic drag of the nacelles of an aircraft equipped with propeller-turbine engines can amount to more than 15 percent of the overall aircraft drag.
In a broad aspect, the present invention provides an air intake system of a propeller-turbine engine with in-line arrangement of a propeller gearbox which combines a reduction of the aerodynamic drag of the nacelle with stable compressor operating conditions of the gas turbine, while being of simple design and form.
It is a particular object of the present invention to provide remedy to the above problem by the features described herein, with further advantages and aspects of the present invention becoming apparent from the description below.
The present invention accordingly provides one or several intake units arranged essentially below a nacelle fairing of the propeller-turbine engine, with a diffuser being related to each intake unit by way of which the flow of air supplied is fed into a rotationally symmetric attenuation chamber which is connected to the compressor inlet of the propeller-turbine engine.
The air intake system according to the present invention is characterized by a variety of merits.
The air intake units according to the present invention, which preferably have the form of so-called NACA air intakes, are arranged essentially under the nacelle fairing. This enables the outer configuration of the nacelle to be optimized and its drag resistance to be reduced.
The in-line arrangement of propeller-turbine engine and propeller gearbox is highly favorable in terms of the space occupied by them, this providing for a smaller overall diameter of the nacelle fairing.
According to the present invention, the number of air intake units may be adjusted to the requirements, which means that between one and five of such intake units may be arranged on the nacelle fairing, for example.
In a particularly preferred form of the present invention, the respective intake unit is provided with a device for the deflection and diversion of the propeller hub boundary layer. Integration of said device for the deflection and diversion of the propeller hub boundary layer in the nacelle surface improves the in-flow characteristics of the intake unit.
In a further, particularly advantageous form, the present invention provides for orientation of the intake units to the nacelle-related speed vector of the propeller flow. This allows the in-flow or by-flow characteristics of the nacelle to be appropriately taken into account to ensure optimum airflow into the intake units.
To reduce losses, it is particularly advantageous to provide each intake unit with a diffuser (elbow-type diffuser) downstream of its inlet cross-section. Such means of shock diffusion improves the inflow of air into the following attenuation chamber. The attenuation chamber ensures a uniform field of flow in the compressor inlet of the gas turbine. This results in stability of the compressor under all flight and load conditions of the propeller-turbine engine.
It is particularly favorable to arrange in the area of the diffuser (elbow-type diffuser) a branch to an inertia-type particle separator, with provision being made for its activation and de-activation.
All intake units of the propeller-turbine engine are connected to the rotationally symmetric attenuation chamber preferably by means of flexible sealing elements. This arrangement provides for a shock diffuser at each outlet of the elbow-type diffuser or each inlet of the attenuation chamber, respectively.
The above-described design, in particular the shock diffuser, separates the intake units from the engine compressor inlet. While this design generally incurs higher pressure losses at the intake of the propeller-turbine engine, the intake configuration so provided appreciably reduces the frontal area of the nacelle of the propeller-turbine engine. The consequential improvement in terms of the aerodynamic drag of the nacelle leads to a significant reduction in fuel consumption. In summary, the smaller frontal area of the nacelle together with lower interference losses provides for reduced fuel consumption, this contributing to the environmental friendliness of the engine and increasing the operating range of the aircraft.
As a further advantage of the present invention, the smaller size of the nacelle fairings provides for a weight reduction of the engine nacelles.
The essential advantage of compressor stability under all conditions of the propeller-turbine engine and all flight conditions has already been addressed above.
Furthermore, engine installation is facilitated by the separation of the nacelle intake (intake units) and the compressor inlet of the engine.
In addition, the attenuation chamber can be provided with secondary inlets for accessories or similar installations, for example oil coolers, cabin-air cooling and/or nacelle venting. Provision is here also made for a reduction of the overall number of secondary inlets.